Gas turbines are used in many fields for driving generators or driven machines. In this case, the energy content of a fuel is utilized for producing a rotational movement of a turbine shaft. To this end, the fuel is combusted in a combustion chamber, wherein compressed air is supplied from an air compressor. The air compressor in this case is customarily constructed as an axial compressor. The operating medium, under high pressure and at high temperature, which is produced in the combustion chamber as a result of the combustion of the fuel, is directed in this case through a turbine unit—which is connected downstream to the combustion chamber—where it expands, performing work. The air compressor, or compressor for short, and the turbine unit are customarily arranged on a common shaft so that the turbine unit drives the compressor during operation.
The combustion chamber of the gas turbine can be constructed as a so-called annular combustion chamber, in which a large number of burners, which are circumferentially arranged around the turbine shaft, open into a common combustion space which is enclosed by a high temperature-resistant surrounding wall. To this end, the combustion chamber is designed as an annular structure in its entirety. In addition to a single combustion chamber, a multiplicity of combustion chambers can also be provided.
For producing the rotational movement of the turbine shaft, in this case a number of turbine rotor blades, which are customarily assembled into blade groups or blade rows, are arranged on this shaft. In this case, a turbine disk, on which the turbine rotor blades are fastened by means of their blade root, is customarily provided for each turbine stage. For flow guiding of the operating medium in the turbine unit, moreover, turbine stator blades, which are connected to the turbine casing and assembled to form stator blade rows, are customarily arranged between adjacent rotor blade rows.
The air compressor of such a gas turbine, in respect to construction, is customarily constructed similarly to the turbine unit and, in a configuration as an axial compressor, comprises a multiplicity of compressor stator blades, which are assembled to form stator blade rows and fastened in each case on a stator blade carrier, and a multiplicity of compressor rotor blades, which are assembled to form rotor blade rows and fastened in each case on a compressor shaft. A rotor blade row and a stator blade row which directly follows it, as seen in the flow direction of the flow medium (in this case air), in this case form a compressor stage. As a rule, a plurality of compressor stages are provided.
The entirety of all the rotating parts of the gas turbine—especially the shaft and the rotor blades—are also referred to as a rotor, and the stationary parts—especially the casing and the stator blades—are also referred to collectively as a stator.
The compressor shaft is customarily assembled from a multiplicity of compressor disks which are arranged one behind the other, as seen in the axial direction, and are held together by means of a tie bolt, for example. In the direction towards the turbine unit, the compressor shaft continues, via a shaft intermediate piece, as the turbine shaft. On its periphery, each of the compressor disks customarily carries the compressor rotor blades of a rotor blade row, which rotor blades, by means of blade roots, for example, are fastened in corresponding fastening grooves of the compressor disk. A compressor disk can also carry a plurality of rotor blade rows.
Annular hollow spaces, also referred to as a cavity in each case in the following text, are customarily provided between two consecutive compressor disks in each case, as a result of which the total weight of the compressor shaft is reduced in comparison to a completely solid type of construction. The last compressor disk, as seen in the flow direction of the flow medium, in a conventional type of construction has a top face or end face which points towards the subsequent turbine unit and together with other components, for example, delimits or at least partially encloses a rear hollow space—also referred to as a rear cavity—which is separated from the flow passage for the flow medium. Such a design is known from EP 1 640 587 B1, for example (compare FIG. 2 there). A plurality of groups of hollow spaces can also be provided in the compressor shaft, wherein, for example, the hollow spaces of a first group lie further on the outside, as seen in the radial direction, whereas the hollow spaces of a second, and, if applicable, of further groups, lie further on the inside.
In the design of such gas turbines, in addition to the achievable power, a particularly high level of efficiency is customarily a design aim. An increase of the efficiency can be achieved in this case, for thermodynamic reasons, basically by an increase of the exit temperature at which working medium discharges from the combustion chamber and flows into the turbine unit. In this case, comparatively high temperatures of the operating medium of 1200° C. and more, for example, are aimed at and also achieved for such gas turbines.
So that such high temperatures or correspondingly high levels of efficiency can be achieved, the air in the compressor should be compressed as intensely as possible. Contingent upon the pressure in the compressor which increases more and more along the flow direction of the gas, the temperature at the compressor exit also increases along with it. The maximum permissible operating temperature of the material of the rear compressor disks is possibly reached in the process.
At present, the maximum permissible operating temperature for available materials represents a limiting boundary condition for the development of gas turbines with regard to the compressor exit temperature. If there is the risk of this limit being exceeded, for example in the case of high ambient temperature, the operating mode of the machine must be throttled. As a result, the potential of the gas turbine cannot be fully utilized.
At the premises of the applicant, technical solutions have been developed, by means of which the last compressor disk, as seen in the flow direction of the flow medium, especially in its top region or end region, is cooled by means of impingement with cooling air. For this, use is made of a cooling-air cooler, for example, which serves essentially for the supply of the front turbine blading with cooled-down cooling air. This cooling air is fed into the rotor through the so-called shaft cover, specifically a shaft cover or casing arranged downstream of the air compressor, as seen in the flow direction of the flow medium. From there, the cooling air then finds its way into the turbine blading. Some of this cooling air is directed from the shaft cover into the cavity downstream of the last compressor disk for cooling the top region of the last compressor disk in the process.
With this form of cooling air feed, it is disadvantageous that the cooling air for the cavity travels a relatively long way through various components of the gas turbine, specifically first through the feed line to the shaft cover and then through the shaft cover itself, which are exposed to circumflow by hot compressor air. As a result, the temperature of the cooling air significantly increases before it reaches the cavity, as a result of which the cooling potential for the last compressor disk is greatly reduced. Furthermore, the last compressor disk is cooled only from one side, and the compressor disks which lie further upstream of it, as seen in the flow direction of the flow medium,—which admittedly are not quite as high as the last compressor disk but are certainly appreciably thermally loaded—are possibly not cooled at all.